Gas turbine engine with two-spool fan and variable vane turbine

ABSTRACT

A gas turbine engine and a method of operating the gas turbine engine according to an exemplary aspect of the present disclosure includes modulating a variable high pressure turbine inlet guide vane of a high pressure spool to performance match a first stage fan section of a low pressure spool and an intermediate stage fan section of an intermediate spool to maintain a generally constant engine inlet flow while varying engine thrust.

BACKGROUND

The present disclosure relates to gas turbine engines, and moreparticularly to a 3 spool variable cycle gas turbine engine.

Variable cycle engines power high performance aircraft over a range ofoperating conditions yet achieve countervailing objectives such as highspecific thrust and low fuel consumption. The variable cycle engineessentially alters a bypass ratio during flight to match varyingrequirements. This facilitates efficient performance over a broad rangeof altitudes and flight conditions to generate high thrust when neededfor high energy maneuvers yet also optimize fuel efficiency for cruiseor loiter conditions.

SUMMARY

A gas turbine engine according to an exemplary aspect of the presentdisclosure includes a low spool along an engine axis with a first stagefan section and a low pressure turbine section, the first stage fansection in communication with a third stream bypass flow path, a secondstream bypass flow path, and a core flow path an intermediates spoolalong an engine axis with a second stage fan section and an intermediatepressure turbine section, the second stage fan section downstream of thefirst stage fan section and in communication with the second streambypass flow path, and a core flow path. A high spool along the engineaxis with a high pressure compressor section and a high pressure turbinesection along the core flow path.

A method of operating a gas turbine engine according to an exemplaryaspect of the present disclosure includes modulating a variable highpressure turbine inlet guide vane of a high pressure spool toperformance match a first stage fan section of a low pressure spool andan intermediate stage fan section of an intermediate spool to maintain agenerally constant engine inlet flow while varying engine thrust.

BRIEF DESCRIPTION OF THE DRAWINGS

Various features will become apparent to those skilled in the art fromthe following detailed description of the disclosed non-limitingembodiment. The drawings that accompany the detailed description can bebriefly described as follows:

FIG. 1 is a general schematic view an exemplary variable cycle two-spoolgas turbine engine according to one non-limiting embodiment;

FIG. 2 is a block diagram of a two-spool fan control algorithm (FCA) foroperation of the variable cycle two-spool gas turbine engine with theacceleration logic flow emphasized; and

FIG. 3 is a block diagram of the two-spool fan control algorithm (FCA)of FIG. 2 for operation of the variable cycle two-spool gas turbineengine with the deceleration logic flow emphasized.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a variable cycle three-spoolbypass turbofan that generally includes a fan section 22 with a firststage fan section 24, an intermediate stage fan section 26, a highpressure compressor section 28, a combustor section 30, a high pressureturbine section 32, an intermediate turbine section 34, a low pressureturbine section 36, a bypass duct section 38 and a nozzle section 40.Additional sections may include an augmentor section 38A among othersystems or features such as a geared architecture which may be locatedin various other engine sections than that shown such as, for example,aft of the LPT. The sections are defined along a central longitudinalengine axis A.

The engine 20 generally includes a low spool 42, an intermediate spool44 and a high spool 46 which rotate about the engine centrallongitudinal axis A relative to an engine case structure 48. It shouldbe appreciated that other architectures, such as a three-spoolarchitecture, will also benefit herefrom.

The engine case structure 48 generally includes an outer case structure50, an intermediate case structure 52 and an inner case structure 54. Itshould be understood that various structures individual or collectivelywithin the engine may define the case structures 50, 52, 54 toessentially define an exoskeleton that supports the spools 42, 44, 46for rotation therein.

The first stage fan section 24 communicates fan flow through a flowcontrol mechanism 56F into a third stream bypass flow path 56 as well asinto a second stream bypass flow path 58, and a core flow path 60. Theflow control mechanism 56F may include various structures such aspneumatic or mechanical operated blocker doors that operate as athrottle point to define a variable area throat and selectively controlflow through the third stream bypass flow path 56 such that a selectivepercentage of flow from the first stage fan section 24 is dividedbetween the third stream bypass flow path 56 and both the second streambypass flow path 58 and core flow path 60. In the disclosed non-limitingembodiment, the flow control mechanism 56F may throttle the flow intothe third stream bypass flow path 56 down to a minimal but non-zeroflow.

The intermediate stage fan section 26 communicates intermediate fan flowinto the second stream bypass flow path 58 and the core flow path 60.The intermediate stage fan section 26 is radially inboard andessentially downstream of the flow control mechanism 56F such that allflow from the intermediate stage fan section 26 is communicated into thesecond stream bypass flow path 58 and the core flow path 60.

The high pressure compressor section 28, the combustor section 30, thehigh pressure turbine section 32, the intermediate turbine section 34,and the low pressure turbine section 36 are in the core flow path 60.These sections are referred to herein as the engine core.

The core airflow is compressed by the first stage fan section 24, theintermediate stage fan section 26, the high pressure compressor section28, mixed and burned with fuel in the combustor section 30, thenexpanded over the high pressure turbine section 32, the intermediateturbine section 34, and the low pressure turbine section 36. Theturbines 32, 34, 36 rotationally drive the respective low spool 42,intermediate spool 44 and the high spool 46 in response to theexpansion.

The third stream bypass flow path 56 is generally defined by the outercase structure 50 and the intermediate case structure 52. The secondstream bypass duct 54 is generally defined by the intermediate casestructure 52 and the inner case structure 54. The core flow path 60 isgenerally defined by the inner case structure 54. The second streambypass flow path 58 is defined radially inward of the third streambypass flow path 56 and the core flow path 60 is radially inward of thecore flow path 60.

The nozzle section 40 may include a third stream exhaust nozzle 62(illustrated schematically) which receives flow from the third streambypass flow path 56 and a mixed flow exhaust nozzle 64 which receives amixed flow from the second stream bypass duct 54 and the core flow path60. It should be understood that various fixed, variable,convergent/divergent, two-dimensional and three-dimensional nozzlesystems may be utilized herewith.

The first stage fan section 24 and the low pressure turbine section 36are coupled by a low shaft 66 to define the low spool 42. In thedisclosed non-limiting embodiment, the first stage fan section 24includes a first stage variable inlet guide vane 68, a first stage fanrotor 70, and a first stage variable stator 72. It should be appreciatedthat various systems may be utilized to activate the variable inletguide vanes and variable stators. It should also be understood thatother fan stage architectures may alternatively or additionally beprovided such as various combinations of a fixed or variable inlet guidevane 68 and a fixed or variable stator 72. The first stage variablestator 72 is upstream of the flow control mechanism 56F.

The intermediate stage fan section 26 and the intermediate pressureturbine section 34 are coupled by an intermediate shaft 74 to define theintermediate spool 44. In the disclosed non-limiting embodiment, theintermediate stage fan section 26 includes an intermediate stagevariable inlet guide vane 76, an intermediate fan rotor 78, and anintermediate stage stator 80. The intermediate stage variable inletguide vane 76 is immediately downstream of the first stage variablestator 72. It should be understood that other fan stage architecturesmay alternatively or additionally be provided such as variouscombinations of a fixed or variable intermediate stage variable inletguide vane 76 and a fixed or variable intermediate stage stator 80.

The high pressure compressor section 28 and the high pressure turbinesection 32 are coupled by a high shaft 82 to define the high spool 46.In the disclosed non-limiting embodiment, the high pressure compressorsection 28 upstream of the combustor section 30 includes a multiple ofstages each with a rotor 84 and vane 86. It should be understood thatthe high pressure compressor section 28 may alternatively oradditionally include other compressor section architectures which, forexample, include additional or fewer stages each with or without variouscombinations of variable or fixed guide vanes. It should also beunderstood that each of the turbine sections 32, 34, 36 mayalternatively or additionally include other turbine architectures which,for example, include additional or fewer stages each with or withoutvarious combinations of variable or fixed guide vanes.

The high pressure turbine section 32 in the disclosed non-limitingembodiment, includes a multiple of stages (two shown) with variable highpressure turbine inlet guide vanes (HPT vanes) 88 between a first stagehigh pressure turbine rotor 90 and a second stage high pressure turbinerotor 92.

The intermediate pressure turbine section 34 in the disclosednon-limiting embodiment, includes a single stage with variableintermediate pressure turbine inlet guide vanes (IPT vanes) 94 upstreamof an intermediate pressure turbine rotor 96. The intermediate pressureturbine section 34 is generally between the high pressure turbinesection 32 and the low pressure turbine section 36 in the core flowpath.

The low pressure turbine section 36 in the disclosed non-limitingembodiment, includes a single stage with variable low pressure turbineinlet guide vanes (LPT vanes) 98 upstream of a low pressure turbinerotor 100. The low pressure turbine section 36 is the last turbinesection within the core flow path 60 and thereby communicates with themixed flow exhaust nozzle 64 which receives a mixed flow from the secondstream bypass duct 54 and the core flow path 60. The augmentor section38A among other systems or features may be located immediatelydownstream of the low pressure turbine section 36.

The first stage variable stator 72 downstream of the variable pitch fanrotor 70 may further include a variable pitch mechanism such as avariable pitch trailing edge flap 72T. The pitch change provided by thevariable pitch first stage fan rotor 70 and or the first stage variablestator 72 facilitates a reduced articulation requirements for thevariable turbine vanes 88, 94, 98 as well as the potential to utilize afixed exhaust nozzle as the third stream exhaust nozzle 62.

Air which enters the first stage fan section 24 is divided between thethird stream bypass flow path 56, the second stream bypass flow path 58,and the core flow path 60 in response to a position of the flow controlmechanism 56F. That is, bypass flow into the third stream bypass flowpath 56 is controlled. The intermediate stage fan section 26 is radiallyinboard and essentially downstream of the flow control mechanism 56Fsuch that essentially all flow from the intermediate stage fan section26 is communicated into the second stream bypass flow path 58 and thecore flow path 60. The variable turbine vanes 88, 94, 98 in therespective turbine sections 32, 34, 36 facilitate performance matchingfor the first stage fan section 24 and the intermediate stage fansection 26 simultaneously to thereby maintain constant engine inlet flowwhile modulating engine thrust.

With reference to FIG. 2, a logic diagram for a two-spool fan controlalgorithm (FCA) 200 is schematically illustrated. The functions of thealgorithm 200 are disclosed in terms of a block diagram. Generally, therectangles represent actions; the parallelograms represent data; and thediamonds represent decision points. It should be understood by thoseskilled in the art with the benefit of this disclosure that thesefunctions may be enacted in either dedicated hardware circuitry orprogrammed software routines capable of execution in a microprocessorbased electronics control embodiment.

A module 202 may be utilized to execute the two-spool fan controlalgorithm (FCA) 200. In one non-limiting embodiment, the module 202 maybe an engine FADEC, a portion of a flight control computer, a portion ofa central vehicle control, an interactive vehicle dynamics simulatorunit or other system. The module typically includes a processor; amemory and an interface. The processor may be any type of knownmicroprocessor having desired performance characteristics. The memorymay be computer readable medium which stores the data and controlalgorithms described herein. The interface facilitates communicationwith the engine 20 as well as other avionics and vehicle systems.

Generally, the first stage fan section 24 is speed matched to theintermediate stage fan section 26 to minimize spillage drag. Thrustchanges are primarily effected with control of the flow and pressureratios through the second stream bypass flow path 58 with theintermediate spool 44.

The rate of fuel flow will be the predominant effect on engine thrustperformance, but the second effect after that is the variable highpressure turbine inlet guide vane 88; the third effect is the variablelow pressure turbine inlet guide vane 98; the fourth effect is thevariable intermediate pressure turbine inlet guide vane 94 and then thefifth effect is the flow control mechanism 56F to control the thirdstream bypass flow path 56. The variable turbine vanes 88, 98, 94thereby facilitate performance matching for first stage fan section 24and the intermediate stage fan section 26 simultaneously to maintainengine inlet flow constant while modulating engine thrust.

Acceleration Scenario

With reference to FIG. 2, under a scenario in which the aircraftairspeed is less than desired, the engine is accelerated as illustratedgenerally by the left side logic of the two-spool fan control algorithm(FCA) 200. Initially, the thrust from the intermediate stage fan section26 is increased through an open modulation of the LPT vanes 98 and aclose modulation of the IPT vanes 94. That is, even with no throttlechange, a resplit of flow to the first stage fan section 24 and theintermediate stage fan section 26 effects a thrust change through themodulation of the LPT vanes 98 and the IPT vanes 94. It should beunderstood that modulation as utilized herein is inclusive but notlimited to any change in any or each of the turbine sections 32, 34, 36.

As inlet air flow has an impact on spillage drag, it is desired tomaximize inlet flow such that if inlet air flow is at maximum, the logiccontinues. If the inlet air flow is below maximum, the third streambypass flow path 56 is modulated toward a more open position through theflow control mechanism 56F.

Then, airspeed is again checked because the third stream bypass flowpath 56 may be modulated toward a more open position through the flowcontrol mechanism 56F such that drag will be relatively decreased, butthrust may be correspondingly reduced. That is, the change in the thirdstream bypass flow path 56 has a relatively smaller impact of thrustcapability, so if the thrust change was not that which is desired, thelogic then changes the throttle and, in this increase thrust scenario,increases the fuel flow rate

Then, there is the rate of change of the fuel flow schedule identifiedas the derivative of the fuel flow rate. Alternatively, Nh dot may beutilized where Nh is high spool speed (rpm) and Nh dot is rev/min/min.So if the desired change is rapid such as a snap acceleration, the HPTvanes 88 are modulated open and the third stream bypass flow path 56 ismodulated toward a more closed position through the flow controlmechanism 56F. To effectuate the snap acceleration, the flow controlmechanism 56F may be rapidly closed as that forces the first stage fansection 24 to a higher pressure ratio. To also effectively accommodatethe snap acceleration and assure the desired HPC surge margin, the HPTvanes 88 are modulated closed if the HPC surge margin is greater thandesired or the HPT vanes 88 are modulate open if the HPC surge margin isless than desired to thereby accommodate the thrust increase.

If the desired airspeed change is relatively gentle, then, depending onwhether or not there is adequate surge margin which is adjusted asdescribed above, the logic basically passes through to the speed checkof the intermediate stage fan section 26.

If the speed of the intermediate stage fan section 26 is increasing asdesired the logic then loops back to the entry point of the two-spoolfan control algorithm (FCA) 200 to repeat the airspeed check.

If the speed of the intermediate stage fan section 26 is increasing aswould be expected in this increase thrust scenario, but at a relativelyslower rate than desired, the thrust from the intermediate stage fansection 26 is further increased through an open modulation of the LPTvanes 98 and a close modulation of the IPT vanes 94. The logic thenloops back to the entry point of the two-spool fan control algorithm(FCA) 200.

Deceleration Scenario

With reference to FIG. 3, under a scenario in which the aircraftairspeed is greater than desired, the engine 20 is decelerated asillustrated by the logic generally along the right side of the two-spoolfan control algorithm (FCA) 200 diagram but is otherwise generallysimilar to the increase thrust scenario described above.

Initially, the thrust from the intermediate stage fan section 26 isdecreased through a close modulation of the LPT vanes 98 and an openmodulation of the IPT vanes 94. As inlet airflow has an impact onspillage drag, it is desired to always attempt to maximize inlet flowsuch that if inlet airflow is at maximum, the logic continues. If theinlet airflow is below maximum, the third stream bypass flow path 56 ismodulated toward a more open position through the flow control mechanism56F.

Then, airspeed is checked. The change in the third stream bypass flowpath 56 has a relatively smaller impact of thrust capability, so if thethrust change was not that which is desired, the logic then changes thethrottle and, in this decrease thrust scenario, decreases the fuel flowrate

Then, there is the rate of change of the fuel flow schedule identifiedas the derivative of the fuel flow rate. So, if the desired change israpid such as a snap deceleration, the third stream bypass flow path 56is modulated toward a more open position through the flow controlmechanism 56F. To effectuate the snap deceleration, the flow controlmechanism 56F may be rapidly opened to force the first stage fan section24 to be quickly at a lower pressure ratio. To also effectivelyaccommodate the snap deceleration and assure the desired HPC surgemargin, the HPT vanes 88 are modulated closed if the HPC section 28surge margin is greater than desired or the HPT vanes 88 are modulateopen if the HPC section 28 surge margin is less than desired to therebyaccommodate the rapid thrust decrease.

If the desired airspeed change is relatively gentle, then, depending onwhether or not there is adequate surge margin which is adjusted asdescribed above, the logic basically passes through to the speed checkof the intermediate stage fan section 26.

If the speed of the intermediate stage fan section 26 is decreasing asdesired the logic then loops back to the entry point of the two-spoolfan control algorithm (FCA) 200 to repeat the airspeed check.

If the speed of the intermediate stage fan section 26 is decreasing aswould be expected in this decrease airspeed scenario, but at arelatively slower rate than desired, the thrust from the intermediatestage fan section 26 is further decreased through a close modulation ofthe LPT vanes 98 and an open modulation of the IPT vanes 94, the logicthen loops back to the entry point.

The deceleration scenario provides no issue for the HPC section 28,however, the LPC section may be subject to surge such that control ofthe flow through the third stream bypass flow path 56 facilitates aneffective change in operating line.

A steady state scenario in which flow through the engine 20 is ineffective balance, the decisions would effectively flow through thecenter of the two-spool fan control algorithm (FCA) 200. That is, thefirst stage fan section 24 is speed matched to the intermediate stagefan section 26 and are in balance, the aircraft is at the desiredairspeed, there is no spillage and there is a desired surge margin onthe HPC and LPC.

The two-spool fan control algorithm (FCA) 200 further utilizes data suchas a speed of the low spool 42, torque on the low spool 42, a speed ofthe intermediate spool 44, the throat area of the flow control mechanism56F into the third stream bypass flow path 56, the throat area of thefirst stage variable stator 72 as well as data such as temperatures andothers.

It should be understood that relative positional terms such as“forward,” “aft,” “upper,” “lower,” “above,” “below,” and the like arewith reference to the engine but should not be considered otherwiselimiting.

It should be understood that like reference numerals identifycorresponding or similar elements throughout the several drawings. Itshould also be understood that although a particular componentarrangement is disclosed in the illustrated embodiment, otherarrangements will benefit herefrom.

Although particular step sequences are shown, described, and claimed, itshould be understood that steps may be performed in any order, separatedor combined unless otherwise indicated and will still benefit from thepresent disclosure.

The foregoing description is exemplary rather than defined by thelimitations within. Various non-limiting embodiments are disclosedherein, however, one of ordinary skill in the art would recognize thatvarious modifications and variations in light of the above teachingswill fall within the scope of the appended claims. It is therefore to beunderstood that within the scope of the appended claims, the disclosuremay be practiced other than as specifically described. For that reasonthe appended claims should be studied to determine true scope andcontent.

What is claimed is:
 1. A gas turbine engine comprising: a low spoolalong an engine axis with a first stage fan section and a low pressureturbine section, said first stage fan section in communication with athird stream bypass flow path, a second stream bypass flow path, and acore flow path; an intermediate spool along an engine axis with a secondstage fan section and an intermediate pressure turbine section, saidsecond stage fan section downstream of said first stage fan section andin communication with said second stream bypass flow path, and said coreflow path; and a high spool along said engine axis with a high pressurecompressor section and a high pressure turbine section along said coreflow path.
 2. The gas turbine engine as recited in claim 1, furthercomprising a flow control mechanism downstream of said first stage fansection, said first stage flow control mechanism operable to throttle aflow in said third stream bypass flow path.
 3. The gas turbine engine asrecited in claim 1, further comprising a combustor section axiallybetween said high pressure turbine section and said high pressurecompressor section.
 4. The gas turbine engine as recited in claim 1,wherein said low pressure turbine section, said intermediate pressureturbine and said high pressure turbine section are along said core flowpath.
 5. The gas turbine engine as recited in claim 4, wherein said lowpressure turbine section includes variable low pressure turbine inletguide vanes.
 6. The gas turbine engine as recited in claim 4, whereinsaid intermediate pressure turbine section includes variableintermediate pressure turbine inlet guide vanes.
 7. The gas turbineengine as recited in claim 4, wherein said high pressure turbine sectionincludes variable high pressure turbine inlet guide vanes.
 8. The gasturbine engine as recited in claim 1, wherein said third stream bypassflow path is radially outboard of said second stream bypass flow path.9. The gas turbine engine as recited in claim 8, wherein said secondstream bypass flow path meets with said core flow path for communicationthrough a convergent/divergent nozzle.
 10. The gas turbine engine asrecited in claim 9, wherein said third stream bypass flow is incommunication with a convergent/divergent nozzle.
 11. A method ofoperating a gas turbine engine comprising: modulating a variable highpressure turbine inlet guide vane of a high pressure spool toperformance match a first stage fan section of a low pressure spool andan intermediate stage fan section of an intermediate spool to maintain agenerally constant engine inlet flow while varying engine thrust.
 12. Amethod as recited in claim 11, further comprising: modulating a variablearea throat of a third stream exhaust nozzle downstream of the firststage fan section.
 13. A method as recited in claim 11, furthercomprising: modulating variable low pressure turbine inlet guide vanesin response to an airspeed; and modulating variable intermediatepressure turbine inlet guide along the intermediate spool in response tothe airspeed.
 14. A method as recited in claim 13, further comprising:modulating open the variable low pressure turbine inlet guide vanes inresponse to the airspeed being less than a desired airspeed; andmodulating closed the variable intermediate pressure turbine inlet guidealong the intermediate spool in response to the airspeed being less thanthe desired airspeed.
 15. A method as recited in claim 13, furthercomprising: modulating closed the variable low pressure turbine inletguide vanes in response to the airspeed being greater than a desiredairspeed; and modulating open the variable intermediate pressure turbineinlet guide along the intermediate spool in response to the airspeedbeing greater than the desired airspeed.
 16. A method as recited inclaim 11, further comprising: modulating variable high pressure turbineinlet guide along the high spool in response to a derivate of a fuelflow rate; and modulating a variable area throat of a third streamexhaust nozzle downstream of the first stage fan section in response toa derivate of a fuel flow rate.
 17. A method as recited in claim 16,further comprising: modulating open the variable high pressure turbineinlet guide along the high spool in response to the derivate of a fuelflow rate increasing; and modulating closed the variable area throat ofa third stream exhaust nozzle downstream of the first stage fan sectionin response to the derivate of the fuel flow rate increasing.
 18. Amethod as recited in claim 16, further comprising: modulating open thevariable area throat of a third stream exhaust nozzle downstream of thefirst stage fan section in response to the derivate of the fuel flowrate decreasing.
 19. A method as recited in claim 11, furthercomprising: modulating variable low pressure turbine inlet guide vanesin response to a speed of the intermediate stage fan section; andmodulating variable intermediate pressure turbine inlet guide along theintermediate spool in response to the speed of the intermediate stagefan section.
 20. A method as recited in claim 11, further comprising:modulating variable low pressure turbine inlet guide vanes in responseto a speed of the intermediate stage fan section changing at a desirerate; and modulating variable intermediate pressure turbine inlet guidealong the intermediate spool in response to the speed of theintermediate stage fan section changing at the desire rate.